The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
Among the engine components, relatively high temperatures are observed in the combustor section such that cooling airflow is provided to meet desired service life requirements. The combustor section typically includes a combustion chamber fondled by an inner and outer wall assembly. Each wall assembly includes a support shell lined with heat shields often referred to as liner panels. In certain combustion architectures, dilution passages direct airflow to condition air within the combustion chamber. In addition to the dilution passages, the shells may have relatively small air impingement passages to direct cooling air to impingement cavities between the support shell and the liner panels. This cooling air exits numerous effusion passages through the liner panels to effusion cool the passages and film cool a hot side of the liner panels to reduce direct exposure to the combustion gases.
With lower emission requirements and higher combustor operational temperatures, effective sealing between the shell and liner panels may be of increased significance. However, relatively large tolerances between the cast liner panels and sheet metal shell may complicate such effective sealing.